The field of the disclosure relates generally to gas turbine engines and, more particularly, to airfoil tip geometry to reduce blade wear in gas turbine engines.
At least some known turbomachines, i.e., gas turbine engines, include a compressor that compresses air through a plurality of rotatable compressor blades enclosed within a compressor casing, and a combustor that ignites a fuel-air mixture to generate combustion gases. The combustion gases are channeled through rotatable turbine blades in a turbine through a hot gas path. Such known turbomachines convert thermal energy of the combustion gas stream to mechanical energy used to generate thrust and/or rotate a turbine shaft to power an aircraft. Output from the turbomachine may also be used to power a machine, such as, an electric generator, a compressor, or a pump.
Under some known operating conditions, rub events occur within the turbomachine, wherein a rotor blade tip contacts or rubs against the surrounding stationary casing inducing radial and tangential loads into a rotor blade airfoil. Generally during rub events, these loads cause the rotor blade to vibrate and deflect causing wear thereto. Excessive tip rub events cause wear to the rotor blade including, but not limited to, loss of blade material, which decreases turbomachine performance.
During tip rub events, the rotor blade is known to lose more material from the tip than the penetration distance into the casing. For example, if the blade tip penetrates the casing 1 mil (25.4 micrometers (μm)) then the blade tip is known to lose as much as 10 mils (254 μm) of material. The thickness of material lost in the blade tip divided by the penetration distance into the casing is known as a rub ratio. In the above example, the rub ratio would be 10:1, or known to have a rub ratio value of 10. Turbomachines with a high rub ratio are known to have decreased performance and decreased service life resulting in higher maintenance costs.